Rocket means for driving a free punch



United States Patent [45] Patented [73] Assignee [54] ROCKET MEANS FOR DRIVING A FREE PUNCH 10 Claims, 8 Drawing Figs.

[52] 0.8. CI 102/49.7 [51] Int. Cl i F42b 13/28 [50] Field of Search 102/49, 70,

70.2, 98; 89/1.7B, 1, 808; 60/35.6R.S.

[56] References Cited UNITED STATES PATENTS 2,924,174 2/1960 McLean 102/49 2,939,275 6/1960 Loedding 102/491 2,945,421 7/1960 Pion 89/1 .73 2,995,01 l 8/1961 Kimmel 60/35.6R.S. 2,997,955 8/1961 Wade et a1 102/49 2,856,820 10/1958 Schmuel et al. 89/1 .7B

3,017,836 l/1962 Guay 102/495 3,069,845 12/1962 Martin et al. l02/49I 3,104,523 9/1963 ODonnell .l 102/49l 3,128,600 4/1964 Oldham 60/356 R.S.

Primary Examiner- Verlin R. Pendegrass Att0rneys-Harry M. Saragovitz, Edward J. Kelly, Herbert Berl and James T. Deaton ABSTRACT: A rocket including: a body; said body including a metal liner with one open end, said liner being of such a thickness as to render said liner capable of withstanding the longitudinal stresses and about half of the circumferential stresses that are exerted by high pressure within said liner, and wrapping material means mounted on said liner so as to render said body capable of withstanding all the circumferential stresses exerted by said high pressure; a solid propellant mounted in said body; igniter means in said body for igniting said solid propellant; a nozzle located at the open end of said body; a spin turbine adapted to be secured to and coaxial of said nozzle along opposed end edges: of said nozzle and turbine, each of said nozzle and turbine having a circumferential groove in the outer periphery thereof and adjacent said opposed end edges; a segmented ring positioned around said opposed end edges and having a portion thereof fitted in each circumferential groove to secure said turbine to said nozzle; and means keying said nozzle and turbine together for simultaneous rotation so long as said segmented ring is positioned around said opposed end edges.

PATENTEI] um 5 I970 SHEET 1 BF 3 Donald V.-B|ack Albert D. Jomtuus Ross T. Rodey John K. Wall,

INVENTORS. 971. M BY J,

PATENTED- DEC] 5 I976 SHEET 2 OF 3 m "mam OO ALTITUDE BURNOUT w or e 4 a o; QHEmLEmE T6753 m om wgtb 638% 9 8 ME;

RANGE (FEET) FIG. 6

Donald V Black Alberr D. Jumtous Ross T. Rudey John K. Wall,

INVENTORS. M 771, BY J, W W cam/L m 7, 90L

SHEET 3 BF 3 V= 7000 FPS FOR 768 77 PENETRATION l2" STEEL TUNGSTEN ESTIMATED llllllllllllllllllll IIIIIIIIIIIIIIYI'IIIII GROSS WT. OF ROCKET.

| I l l Y I I 1.0 SLUG DENSITY LB/IN3 1o lllllll llll VELOCITY K I000 FT/SEC) PEG. 7

km E fy m m fiHu OO v v w db l dawn m m mDARJ b :m 9 8 7 S m N KT CE 4 IM A m B R u C OE N 13 MR mw Y n U A Q 2 B O m M R A h u E h n h n w W98? 6 5 4 3 2 FIG. 8

ROCKET MEANS FOR DRIVING A FREE PUNCH This application is a division of application Ser. No. l96,854,f1led May 18,1962.

The present invention relates to a device-for punching holes in massive structures such as heavy armor, spaced or continuous, having a total thickness of 6 to 12 inches for example and relates more particularly to a missile drive system and the components thereof. Such armor is found on naval craft and on land vehicles such as tanks. Many devices have been tried for disabling tanks, such as land mines, artillery and bazookas, for examples. Resistance to or the foiling of such attack devices has been built into tanks as fast as such devices have appeared.

There is need for a device which will punch a hole in any armor which a tank or ship could conceivably carry. There is need for a punch which will not be foiled as by false coverings which either deflect or trigger premature action, or foiled by layers of spaced armor plate. And there is need for a punch which can be launched from and be effective at distances of a quarter to about two miles from the target. Further, the launching of the punch must be achieved by, preferably, one man aiming and firing equipment which can be easily carried by him. Visual sighting of the launcher must be simple and quick. The weight of a complete launcher, power plant, and punch should not exceed twenty pounds if the device is to fulfill such requirements. This performance ability with weight limitations is not found in the prior art. Also there is need for a punch for use in attacking concrete and rock. The concrete may be either reinforced with steel rods or steel aggregate, or unreinforced with such materials. In fact, there is need for a punch which will effect penetration of a wide variety of materials for a wide variety of purposes. These purposes occur both under military and under civilian conditions.

It is imperative to the utility of a missile that it have a very high probability of a first-round hit. Until very recently, this requirement could not be met by high-velocity guns, except, of course, at extremely short ranges, pointblank, a condition where unguided low velocity rockets carrying shaped charges are fairly adequate. However, guns large enough to be lethal to a heavy tank are themselves very heavy and expensive. As a result, adequate antitank gun support is rarely available to troops under tank attack. Recently, small guided missiles of low velocity and using shaped charges have become available. These missiles have the advantage over guns of only requiring light and portable launchers, but, primarily, due to their guidance means, they are relatively unreliable and expensive.

The present invention relates to a rocket and punch which fills this need for an antitank weapon, a weapon that can be handled and fired by a single man, that has an effective range of a it to 2 miles, and that has a high probability of first-round hit and kill. The present invention, also, relates to a free punch which can be launched remotely from a target, will have a high degree of accuracy as to hitting the target just where desired, and which will penetrate through or deeply into any material.

Thus, it will be seen that it is an object of the present invention to devise a system for punching holes in massive objects such as thick armor plate by means of a punch which is launched remotely from the target and which systems components are very light in weight, and which punch can be put on target by a simple and direct sight because of a very short flight time.

Further objects of the present invention result from fundamental aspects of the present device. If the weight of the device is to be within the given limit, the punch weight must be close to or under one pound, the power plant must be under twelve pounds, and the cartridge-launcher must not exceed pounds. The only.manner in which a free one pound punch can penetrate twelve inches of armor is for the punch to have a density, a slenderness, and a velocity that will give a high enough impact momentum per unit area of punch cross section or unit area of punch and target contact to be effective.

This means a critical proportioning of, weight, power, and

dimensions. The weight and dimensions of a launcher and a rocket motor to drive a punch are critical if mobility is to be achieved with needed punch weight and velocity. Thus, the power available is critical. The power available dictates the weight of the punch, and this dictates the density and slenderness of the punch necessary to achieve the unit impact pressures required to effect penetration. The prior art does not teach such proportioning or how to achieve such unit pressures. Further, the art does not teach how to deliver such a free punch in proper aspect, so that its longitudinal axis is tangent to its flight path.

Thus, it becomes a further object of the invention to provide a punch and rocket drive therefor which may be transported and launched by a single person such as a foot soldier, and which punch and rocket, and punch, rocket, and launcher are so proportioned to each other and to their components that such proportioning is critical and critical to the resulting punch function.

Further objects relating to the carrying out of the primary objects are the design of a rocket case which is light enough and strong enough so that the primary objects may be achieved, and which case is strong enough so that the ratio of casing weight to grain weight is smallerthan formerly thought possible; and the proportion of weight to punch velocity is more favorable than formerly thought possible; the design of a turbine which will spin the rocket and punch sufficiently to insure a highly accurate flight, and which will drop off at the end of the launching; the design of a connector between rocket casing and punch which will hold them together in flight,

which allows easy punch exchangeability, andwhich permits and induces desirable energy exchange between rocket and punch; the design of sabots which will drop off at launch and which will minimize tipoff; the design of a rocket and launch system so that for a given range there: will be self-compensation in the system for crosswinds; and the design of a rocket so that aiming errors will be practically eliminated due to the velocity of its flight, its time of flight being of the order of one second.

The aforementioned objects and others herein apparent are achieved by a rocket having in its nose cone a cylindrical punch, or slug, with its axis coaxial of the cone and the body of the rocket; which slug is slender and weighs about a pound, and is made of a dense material such, as tungsten or an alloy thereof, or depleted uranium-238, or an alloy thereof, and which rocket has a grossweight of about ten pounds, ready to launch but exclusive of its launcher, with a capability of achieving a speed of 8,000 feet per second in 0.6 seconds; and with a spin turbine, which drops off at launch, to give the rocket accuracy in flight. Such a rocket may be launched in a manner similar to a bazoolca and with similar equipment but will have an effective range of about l,200 to 10,000 feet for heavy armor penetration. Such a rocket will have a very flat trajectory and may be sighted with a conventional optical system with little lead and elevation.

Hereinafter, there is described in detail a rocket device conforming to the above outline and capabilities, which achieves and will achieve the aforementioned objects, and which is illustrated in the drawings herewith in which:

FIG. 1 is an elevational view of a rocket embodying the present invention, which rocket is shown in its launching tube shown in longitudinal section, and which rocket is shown with its firing means ready for firing.

FIG. 2 is a longitudinal sectional view of the rocket and launching tube of FIG. l but without the firing mechanism, and with parts of the rocket casing in full view.

H6. 3 is a sectional view on the line 33 of FIG. 1.

FIG. 4 is a detailed isometric view of a portion of the firing means and of a portion of the rocket case to which such means are to be attached.

FIG. 5 is a detailed isometric exploded view of a spin turbine and the locking sections which aid in securing the turbine to the back end of the rocket, shown fragmentally.

FIG. 6 is a graph of various operating characteristics of the rocket against rocket travel distance.

FIG. 7 is a graph of estimated gross weights of rockets against rocket velocities, and punch densities and slenderness.

THE ROCKET A rocket assembly or vehicle 11 embodying the present invention is shown in FIG. 2 with parts thereof cut away to illustrate the internal construction thereof. There are three main sections of the rocket; a flared nozzle section 12, a fuel containing body section 13, and a nose section 14 which contains a punch or punch member 16. The body or fuel case 13 must be designed to withstand pressures of 3,500 to 4,000 p.s.i. during burning of the fuel contained therein. In closed end thin wall pressure vessels, the longitudinal stresses are one-half the circumferential stresses. This means that in a pressure cylinder made of homogenous material, the material is understressed longitudinally. The present construction materially reduces the weight of the rocket body by using an aluminum liner 17 which is only thick enough to carry the longitudinal stresses and about half of the circumferential stresses. The-other half of the circumferential stresses are carried by a circumferential wrapping 18, or winding, of high strength fiberglass cords which are stabilized, held in place and protected against abrasion, by impregnation of the wrap with an epoxy resin and curing agent. The tension of the wrap is such that the aluminum liner, under pressure, is allowed to develop its full tensile strength circumferentially. Another function of the aluminum liner is that of ceiling the wrap formed by the glass cords and resin against radial gas leakage under pressure. A still further aspect of the use of an aluminum liner is that it may be anodized so as to surface it with an aluminum oxide coating which is good refractory and is resistant to the temperatures and abrasion to which the inside surfaces of the liner are subjected during the burning of the contained solid fuel 19.

A rocket designed in accordance with the present disclosure hasan overall length of 41.20 inches, a body outside diameter of 2.60 inches, a nose cone length of 10.0 inches and a nonle length of 5.70 inches with an outside diameter of 3.56 inches. With these dimensions, if the liner thickness'is 0.032 inch and the wrap thickness is 0.017 inch with respective densities of 0.1 and 0.07 lb./in. there willbe a saving of almost a pound in weight of the body as compared with the use of aluminum only (double the thickness of the liner) in the body.

The liner material is carried aft of the body section and flared to provide a covering 21 for a ceramic nozzle 22. The outside of the nozzle sleeve is in the general shape of a truncated cone fitted and secured to the inside of the cover 21.

' The after edge portion of the nozzle extends to about the edge of the covering 21 which is adapted to butt a portion of a spin turbine. Adjacent this edge, the cover 21 is formed with an external peripheral groove 23 that cooperates with such turbine. The throat 24 of the nozzle is in its forward part, and the inside of the nozzle is flared outwardly both forwardly and rearwardly from its throat. The use of ceramic protects the nozzle 22 from the high heat-loads imposed by the exhaust.

The forward end, the nose end, of the body is closed by a heavy steel or aluminum nose disc 26, skirted and domed, that is convex forwardly. The after outside portion 27 of the skirt is formed parallel to and contiguous with the inside of the liner .47 at its forward end, and thereat the liner and the disc are welded together. The forward portion 28 of the skirt is threaded to receive thereover and retain the base portion of a nose cover cone 34 which provides fairing for the punch 16. The voids inside of the nose cone may be filled by a foamedin-place plastic 35 to reinforce the cone so that it may be formed of light weight material.

The punch is, for example, a solid cylinder 7.77 inches long and 0.52 inch thick to give a slendemess ratio of 15. It is made of tungsten or uranium or an alloy of either material. The density will be of the order of 0.53 to 0.66 lb./in. The aft end of the punch is threaded to screw into a socket 36 located cenwell known to rocket designers. It is sufficient to state that the trally of the nose, disc 26 so as to hold the punch with longituv dinal axis coaxial of the rocket. The threading of the punch to the disc 26 makes for easy assembly and for change of punches depending on the type of target.

The above details describe the net, or empty, rocket. Added to this to give the gross, or full, rocket is the solid propellant fuel 19, or grain, and its igniter 38 with its fuze pig tail 39. The igniter 38 is nozzle-throat mounted for ease of installation and to serve as a weather seal at the throat 24 for the grain under ready conditions. It is full grain-length to achieve full ignition with low time delay and is center supported by consumable foam annuli 40. There may, also, be included in the gross of Y the rocket, a spin turbine 41 that is secured aft of the nozzle 22 but drops off after launching of the rocket. Many considerations go into the selection of a suitable fuel, and these will not be discussed in detail here as such considerations are fuel will be of the solid type such as Omax S-2b which is cast in place in the rocket body. Omax S-2b is a product of Olin Mathison Chemical Corporation. The grain 19 will have a cross section as shown in FIG. 3. The grain should have a burning time of about 0.3 to 0.5 seconds at to -40 F. and develop pressures between 3,500 and 2,000 p.s.i. at substantially constant pressure during burning.

LAUNCH AIDS The rocket may be shipped in a container which, also, serves as a launching tube, and such will be here described. The previously described rocket is shown in such a tube in both FIGS. 1 and 2 which are elevational views with parts sectioned and broken away for clarity of detail. This launching tube 42, shown in longitudinal section in both views, may be constructed in the same manner as the body 13 of the rocket by the use of a anodized aluminum tubular cylinder for the liner and a stabilized wrap of fiberglass cords and cured resins, but it is here shown as a plain aluminum tube. Each of the open ends of the tube is selectively closable by means of a friction fitting closure 46 shown only in FIG. 2. The length of the tube 42 is slightly more than that of the rocket. The after edge of the launch tube, the edge adjacent the rocket nozzle, is formed with a small slot 47 by means of which there is, at the time of firing, secured inside the tube a percussion type firing cap 48 that'has attached thereto one end of the detonator pig moisture and dirt. This tube may serve as a cartridge case or as a launching tube when it is auxiliary equipped with trigger mechanism, sight, and holding or support means.

The illustrated auxiliary launching equipment of FIG. 1 is only that necessary for a shoulder supported launch. This equipment may be secured to the tube directly, or to spaced apart straps 49, each circumferential of the tube 42, as illustrated. A shoulder rest 50 is secured to the straps 49 just forward of the center of gravity of the assembly, and forward of the rest is a pistol grip shaped hand hold 51. The pistol grip 51 carries a trigger 54 that is connected, or extended, by covered cable 56, Bowden wire, to a double acting trigger mechanism and firing pin 57 actuated thereby and contained in a firing block 58 which is placed in opposition to the firing cap 48 so that first actuation of the trigger will cause the pin 57 to cock, and a second actuation will release the pin to impinge the cap 48 to cause its detonation and initiation of the burning of the pig tail 39 to the rocket fuel igniter 38. Firing pin actuation is an old art, and the details of the present actuation are not shown. The firing block 58 is releasably secured in a slot 59 formed in the after edge of the tube adjacent the firing cap slot 47, and the cap and firing block are so aligned when in their respective slots 47 and 59 that upon actuation of the firing pin, it will strike and detonate the cap. This arrangement permits easy attachment and disengagement of the cap and firing block to and from the launching tube when the rear closure 46 is-removed.

Another piece of equipment used in launching the rocket is a sabot 61. The one herein illustrated and described in full length one made in three sections 62, 63, and 64 or petals which extends longitudinally of the rocket and circumjacent the body thereof between body and launcher tube 42 and a portion of the nose cone 34. The rocket rides on and is guided by the sabot as it travels through the tube in its launch. The sabot petals separate and fall from the rocket as it, and they leave the tube. The sabot is made of balsa wood or other light material.

The last piece of equipment needed for the launching of the rocket is the spin turbine 41. This turbine comprises a one inch long cylindrical ring 66 having the same inside diameter as the inside of the after edge of the nozzle 22, with a plurality of blades 67, about twenty, secured inside of and to the ring. Each blade is set at thirty degrees with the center line of the rocket and protrudes 74 inch centripetally of the ring. The forward edge of the ring is flanged outwardly and butts the rear edge of the nozzle cover 21. Also, the butting flange of the ring forms an exterior groove 68 similar to, close to, and in axial alignment with the groove 23 in the nozzle cover 21. Also, the rear edge of the nozzle is formed with three equally circumferentially spaced lugs 69 that extend forward through notches in the butting edge flange of the noule. Three pieces of locking ring sections 71, each having a U-shaped cross section, encompass these circumferential edges, one leg of the U- shape being in one groove 23 and the other leg in the other groove 68. Each ring section is between two of the lugs 69. These ring sections 71 are retained in place by the inside of the launch tube 42 when the rocket is therein, and the rings fall away or are forced by centrifugal force from the rocket upon the rocket and turbine leaving the launch tube. The lugs 69 serve to prevent rotation of the turbine with respect to the rocket and to center the turbine coaxially thereof by the lugs contacting the outside of the nozzle. The turbine is formed of cast stainless steel. The turbine blades are not shown in FIG. 1.

LAUNCH AND FLIGHT The design and fueling of the rocket are such that in flight it will have a very flat trajectory whose apogee is less than 4 feet for target ranges of 1.25 miles. For this range, small changes in missile speed have a negligible effect on the impact point. For example, a one per cent change in the burnout speed results in only a 6 inch vertical change in the impact point. If probable deviations from the nominal of drag, air density, wind velocity, impulse, and weight are all present, the probable vertical error would be less than a foot for such range. While horizontal and vertical dispersion is important for a rocket of this type, due to the rockets small size, light weight, long and slender shape, rotation, and high acceleration, it is possible to design it in such a way as to minimize the sources of such errors. Because of the rocket's small size and weight, it can be launched from a light and mobile tube and aimed by direct sight. From firing, the rocket will travel the first mile in 0.95 seconds so that at this range, the gunner need lead a tank traveling at thirty miles per hour by only about half the length of the tank.

The rocket is launched from the tube 42 with a close fit between the rocket and tube being maintained by the use of the sabot assembly 61. The rocket is guided for about 3 feet of travel before the turbine 41 passes out of the tube. During this periodQthe rocket is accelerated to a roll rate of 800 radians per second by the reaction type spin turbine 41 attached to the rear of the nozzle 22. After leaving the launch tube, the spin turbine 41 and the sabot petals, 62, 63, and 64 are separated from the rocket by the action of the centrifugal force and rocket accelaration. When the turbine leaves the tube, the inner wall of the tube no longer holds the locking sections 71 in place to retain the turbine on the open end of the nozzle cover 21. The rocket leaves the launcher with a forward velocity of approximately 220 feet per second and an acceleration of approximately 360 gs. Over the range for which this rocket is designed to operate, the effect of a crosswind is target. The configuration of the nozzle cover 21 is such that the rocket will have this required stability, that its yaw will be minimal, and that the rocket axis and the axis of the punch 16 will be tangential to the line of flight so that upon impact the punch will be properly presented to the target.

The curves of FIG. 6 depict the speed 72; its time of flight, curve 73; and its altitude during flight, curve 74. These curves are illustrative of the short time between firing and impact on the target. The altitude curve shows that small errors in aiming will have little effect on the accuracy and probability of a hit withinthe intended range of use. The short time to maximum casing pressure and velocity means that the device is effective at short ranges.

PENETRATION OF TARGET The curves of FIG. 7 are plotted against an ordinate representing estimated, or design, gross weights of rockets. The velocity constant of 7,000 f.p.s., for curves 76 and 77, is close to the minimal velocity needed to obtain the desired flat trajectory and the desired time between firing and impact to reduce errors caused by movement of the target so that direct sighting of the target may be used. This minimal velocity is about 5,000 feet per second. It is, also, the minimal velocity which will give, over the indicated range, the desired minimal impact pressures. While penetration; is a function of slenderness and density, it is, also, a direct: function of velocity but increases in velocity result in a geometric increase in the total weight of the rocket as is evident from curve 75 of FIG. 7, a plot of ordinal estimated gross weights against velocity. Thus, the velocity must be kept to that minimum which will give the needed penetration if the desired low rocket weight is to be achieved. According to curve 75, this minimum is about 5,000 f.p.s. Curve 76 depicts rocket weight versus density of the punch load as the abscissa where the punch has a slenderness ratio of eight, an expected penetration of 12 inches, and a velocity of 7,000 feet per second. This curve indicates that, with these constants, to be able to achieve a light weight rocket that can be fired by a single man, punches of a density greater than 0.5 pounds per cubic inch must be used. Curve 77 shows the requirement for slenderness ratios of about ten and much greater. The abscissa of this curve are values of the ratio Aberdeen Proving Grounds as reported in Ordnance Corp.-

Pamphlet, ORDP 20-245, May 1957, page 2-ll9, and the results of these shots at 60 obliquity to the target normal are shown in FIG. 8 by points A and B which are plots of ballistic limit against penetration (normal to face) over core diameter, or punch diameter. These punches were formed from steel. The other points, C to I, inclusive, on this graph represent the test data from shots made by the above BRL at applicants request to check the concept of the herein disclosed invention. The target consisted of two parallel plates of 2 inch and 4 inch thick homogeneous armor, separated by a 6 inch air gap. These shots were, also, made at sixty degrees to the armor normal but were made at higher velocities and with punches of greater density and slenderness. From this figure, it will be seen that there is poor or no correlation between the prior art and the present disclosure. Penetrations: have been achieved which were not predictable for increased penetration should lie on the lines through A and B and there above. On these of the rocket, curve lines, penetration is a function of velocity. The test data for points C to l, inclusive, is given in the following table:

Pene- P 005 60 Length, Diameter, Density. tration, Velocity, Weight, Shot in i 's inches L/l) ll)./in. inches 1) ft./see. lb.

0.. 6.0 .600 10 .526 4.0 3.3 5,359 .84 I)- 5.7 .570 10 .611 7.2 6.3 T, 500 .83 E. 5. 6 500 10 636 7. 7 6. 7, 502 814 F.. T. 47 530 14 '26 7.8 7. 4 7, 519 896 G... T. 47 530 14 .526 8. 1 7. 6 7, 538 910 ll 7. 55 .514 14.6 .600 0.5 9.2 7,491 .858 1.. 7.77 .517 15 .526 0. 8 9. 5 7, 600 .834

In the above shots, the penetration of both plates was in the line of impact. There was no deflection at either the first or second faces. Thus, spaced armor does not act as a foil to the punch of the present invention. Also, as these punches do not employ a shaped charge, they are not foiled by false coverings to a target. Further, it is believed that in actual use, there will be some energy transfer from the rocket case and the punch mounting disc 26 to the punch, and from the case to the disc and then to the punch, upon impact of the punch on the target.

In the explosive field, penetration of armor by means of cavity-lined shaped charges is considered to be probably the most efficacious method, yet, penetration achieved as indicated hereinbefore approximate those for an idealized jet where the velocity is above 20,000 feet per second. The penetrations depicted in FIG. 8 by the points C to l, inclusive, are within seven to twenty per cent of those which would be obtained by an idealized jet where the penetration is a function of the punch length times the square root of the ratio of punch and target densities. See: The science of High Explosives by Melvin A. Cook, page 252, published 1958 by Reinhold Publishing Corporation, New York, New York. The above idealized formula for jet penetration, when once appreciated as being applicable to a high velocity punch, indicates the need for a punch that has a greater density than the target and, also, it indicates the probable length of punch that must'be used to penetrate a given piece of armor. However, the above reference to shaped charge phenomena is not a statement of the theory of penetration found in the present invention. It is only a comparison of results. Further, shaped charge phenomena does not teach the obtaining of such penetrations at the velocities nor by the means disclosed herein.

Given the structures set forth herein which it may be desired to penetrate, such as heavy armor, we hereby disclose a method and a device for the practice of such method whereby a punch may be launched at distances up to 2 miles from such structures by a single person using a device weighing less than twenty pounds, and which can be directly sighted and fired, and which punch will penetrate such structyres.

We claim:

1. A rocket including: a body; said body including a metal liner with one open end, said liner being of such a thickness as to render said liner capable of withstanding the longitudinal stresses and about half of the circumferential stresses that are exerted by high pressure within said liner, and wrapping material means mounted on said liner so as to render said body capable of withstanding all the circumferential stresses exerted by said high pressure; a solid propellant mounted in said body; and igniter means in said body for igniting said solid propellant; a nozzle located at the open end of said body; a spin turbine adapted to be secured to and coaxial of said nozzle along opposed end edges of said noule and turbine, each of said nozzle and turbine having a circumferential groove in the outer periphery thereof and adjacent said opposed end edges; a segmented ring positioned around said opposed end edges and having a portion thereof fitted in each circumferential groove to secure said turbine to said nozzle; and

means keying said nozzle and turbine to ether for sim ul taneous rotation so long as said scgmente ring 15 positioned around said opposed end edges.

2. A rocket as set forth in claim 1, wherein said liner is made of aluminum and the inner surface of said liner is anodized to render the inner surface resistant to temperatures and abrasion to which the inner surface of the liner is subjected.

3. A rocket as set forth in claim 1, wherein said liner is made of aluminum and said wrapping material means includes high strength fiberglass cords which are stabilized, held in place and protected against abrasion, by impregnation of the wrap with an epoxy resin and curing agent.

4. A rocket as set forth in claim 3, wherein said wrapping material means is so tensioned about the aluminum liner that the aluminum liner, under pressure, is allowed to develop its full tensile strength circumferentially.

5. A rocket including: a body; said body including a metal liner with one open end, said liner being of such a thickness as to render said liner capable of withstanding the longitudinal stresses and about half of the circumferential stresses that are exerted by high pressure within said liner, and wrapping material means mounted on said liner so as to render said body capable of withstanding all the circumferential stresses exerted by said high pressure; a solid propellant mounted in said body; and igniter means in said body for igniting said solid propellant, said igniter means including an igniter extending the full length of said solid propellant, consumable foam annuli supporting the center portion of said igniter, and weather seal means at one end of the igniter and mounted in a nozzlethroat of said rocket.

6. A rocket including: a hollow elongated body open at one end and closed at the other; a solid propellant mounted in said body; a drive nozzle mounted relative to said body at said open end; an igniter extending substantially the length of said propellant and mounted in the throat of said nozzle to form a weather seal at said throat; consumable annuli means at spaced intervals along said igniter supporting said igniter in the center of said propellant; and a spin turbine secured to said nozzle along opposed end edges of said nozzle and turbine, said nozzle and turbine being secured together by a segmented ring positioned around opposed end edges of a circumferential groove adjacent said opposed ends of said nozzle and turbine, by portions of the segmented ring that are fitted in the circumferential grooves, and by keying means that keys said nozzle and turbine together for simultaneous rotation so long as said segmented ring is positioned around said opposed end edges.

7. A rocket having a drive nozzle, a spin turbine adapted to be secured to and coaxial of said nozzle along opposed end edges of said nozzle and turbine, each of said nozzle and turbine having a circumferential groove in the outer periphery thereof and adjacent said opposed end edges, a segmented ring positioned around said opposed end edges and having a portion thereof fitted in each circumferential groove to secure said turbine to said noule, and means interconnecting said nozzle and turbine together for simultaneous rotation so long as said segmented ring is positioned around said opposed end edges.

8. A rocket as set forth in claim 7, wherein said turbine is propelled by hot gases communicated through said nozzle, and said nozzle has an outer radially extending surface that extends outwardly relative to other rocket body structure to said spin-turbine.

9., A rocket as set forth in claim 7, wherein said turbine has a central opening with turbine blades secured therein.

10. A rocket as set forth in claim 7, wherein said segmented ring is adapted to be removed from said nozzle and turbine by centrifugal force during the flight of the rocket. 

